One-piece closed-shape structure and method of forming same

ABSTRACT

The present invention relates to a one-piece closed-shape structure and a method for manufacturing a one-piece closed-shape structure. In particular, the present invention relates to a one-piece fuselage and a method for manufacturing a one-piece fuselage. One embodiment of the method of the invention involves the use of molding technology, tooling technology, the integration of the molding and tooling technology, and fiber placement to manufacture a one-piece closed shape structure.

I. CROSS-REFERENCE TO RELATED APPLICATION

This application claims the benefit of U.S. Provisional Application No.60/248,190, filed Nov. 15, 2000 by Alan H. Anderson, Kathlene K. Bowman,and Paul D. Teufel and titled ONE-PIECE CLOSED-SHAPE STRUCTURE ANDMETHOD OF FORMING SAME, the disclosure of which is expresslyincorporated herein by reference.

II. BACKGROUND OF THE INVENTION

A. Field of the Invention

The present invention relates to a one-piece closed-shape structure anda method for manufacturing a one-piece closed-shape structure. Inparticular, the present invention relates to a one-piece fuselage and amethod for manufacturing a one-piece fuselage.

B. Background of the Invention

Since the 1940's and 1950's, aircraft have been manufactured fromlightweight metals, primarily aluminum. More recently, compositematerials (such as fiber reinforced plastics) have been used tomanufacture some aircraft. The manufacture of such aircraft include themanufacture of the fuselage (the central body of the aircraft), thewings, and the various other components of the aircraft.

In the manufacture of an aircraft fuselage with metals or composites,the typical manufacturing process involves the combination of severalpieces that are individually manufactured and then bonded together toform the fuselage. These multiple steps have many disadvantages,including both high cost and significant time.

The creation of a single-piece fuselage would provide many advantagesover fuselages manufactured from the combination of multiple parts.These advantages potentially include lower cost, lighter weight,improved integration, safety, improved performance, noise reduction,improved aerodynamics, and styling flexibility.

As for lower cost, a one-piece fuselage is less costly to fabricate,because there is only one part to manufacture, and there are nofasteners. Thus, the one-piece design saves money in both thefabrication stage and in combination stage. In addition, the work areasneeded at a manufacturing facility are less for a one-piece design,because multiple parts require dramatically more workspace areas.

As for lighter weight, because there are fewer parts to a one-piecefuselage, and because there are fewer fasteners, a one-piece fuselage islighter than a fuselage created from multiple parts. The lighter theaircraft, the more carrying capacity that the aircraft will have, whichis a substantial benefit.

As for improved integration, a one-piece fuselage is easier to integratewith the other components of the aircraft, such as the tail cone, thewings, and the other parts of the aircraft. Additionally, the interiorof a one-piece fuselage would also be easier to integrate, because thereis only one form that must be properly fitted. Moreover, problems withintegration of multiple parts (such as dimension variation and otherfabrication problems) would be completely eliminated in a one-piecedesign.

As for safety, a one-piece fuselage offers structural advantages over afuselage fabricated from multiple parts. In the initial fabrication ofthe one-piece fuselage, the structure may be designed with safetyimprovements (such as strengthened areas, etc.). Additionally, becausethe one-piece fuselage does not have most of the fasteners necessary forcombining the multiple parts, the one-piece design is more structurallysound, which provides increased passenger safety. Also, a one-piecefuselage is more crashworthy. A one-piece fuselage provides theadvantages of an integrated structure, which has numerouscrashworthiness benefits.

As for improved performance, there are both objective and subjectiveimprovements. For objective improvements, there is of course theimproved aerodynamics, which results in greater speed. For subjectiveimprovement, there is the noise reduction, which results in a morecomfortable ride. In some way, all of the advantages of the one-piecefuselage play a role in improved performance.

As for noise reduction, because a one-piece fuselage would result inimproved aerodynamics, a further benefit would be a diminution of airdisruption, which results in noise reduction. Any increase in thesmoothness of an aircraft has the benefit of noise reduction. Thus, tothe extent that the creation of a one-piece fuselage results in theimprovement of aerodynamics, there is a reciprocal decrease in noise.

As for improved aerodynamics, a one-piece fuselage inherently is moreaerodynamic than a fuselage created from the combination of multipleparts. This improvement in aerodynamics would result from the absence ofseams or joints as well as the absence of rivets or other externalfasteners. In modern aircraft, seams and joints between the combinedparts increase drag and thus diminish aerodynamics. By omitting theseams and joints in a one-piece fuselage, aerodynamics would beimproved. Also, in modern aircraft, the external fasteners for flangesand other structure internal to the fuselage also increase drag anddiminish aerodynamics. A one-piece fuselage would omit most fastenersand would thus improve aerodynamics.

As for styling flexibility, the capability to create a one-piecefuselage would provide more opportunities for aircraft design. Becausemultiple parts are not combined to create the fuselage, unique shapesmay be possible, that were previously difficult to achieve. By improvingthe design and styling of the aircraft with a one-piece fuselage, itwould thus be possible to create a more attractive aircraft for themarket.

Therefore, it is desirable to provide a one-piece fuselage.

For a one-piece fuselage, either metal or composite materials may beused. Metal has more disadvantages, due to the inability to fabricateall components of the fuselage in a single step. Composite materials arethus more advantageous for the fabrication of a one-piece fuselage,because composite materials may be fabricated simultaneously.

Therefore, it is further desirable to provide a one-piece fuselagemanufactured from composite materials.

Methods and structures in accordance with the invention provide for aone-piece structure manufactured from composite materials, including aone-piece fuselage. One embodiment includes manufacturing a one-piecefuselage by filament winding. Other embodiments for manufacturing aone-piece fuselage may also be used.

III. SUMMARY OF THE INVENTION

Methods and structures consistent with the present invention mayovercome the shortcomings of conventional systems by providing aone-piece closed shape structure manufactured by composite materials.Additional objects and advantages of the invention will be set forth inpart in the description, which follows, and in part will be obvious fromthe description, or may be learned by practice of the invention. Theobjects of the invention will be realized and attained by means of theelements and combination particularly pointed out in the appendedclaims.

In accordance with an embodiment of the present invention, a method ofmanufacturing a one-piece closed-shape structure using a mandrelcomprises: preparing the mandrel, wherein the mandrel comprises a bagand an armature; applying a frame mandrel to the mandrel to form a framefor the structure; filling the mandrel and the frame mandrel with media;applying a curable resign to a fiber; applying the fiber over themandrel and frame mandrel to form the structure; curing the structure;removing the media from the mandrel and frame mandrel; and extractingthe mandrel and frame mandrel from the structure.

In accordance with another embodiment of the present invention,preparing further comprises: placing the armature through the bag andconforming the shape of the bag to a desired shape of the structure.This embodiment may also include sealing the bag; placing the armatureand the bag in a form tool; and conforming the shape of the bag to theform tool. Further, this implementation may include filling a spacebetween the armature and the bag with air and creating a vacuum betweenthe form tool and the bag to force the bag to conform to the shape ofthe form tool.

In accordance with another embodiment of the present invention, applyinga frame mandrel further comprises applying the frame ply to an exteriorof the bag and applying the frame mandrel over the frame ply.

In accordance with another embodiment of the present invention, fillingfurther comprises compacting the media. In this embodiment, compactingmay further comprise vibrating the mandrel and frame mandrel to aidcompaction.

In accordance with another embodiment of the present invention, applyingthe fiber comprises winding the fiber over the mandrel and frame mandrelto form the structure. In this embodiment, winding may further includeplacing a first winding aid on the bag; winding the fiber over the firstwinding aid, the frame mandrel, and the mandrel to form an inner skin;cutting the inner skin to remove the first winding aids; placing asecond winding aid on the inner skin; winding the fiber over the secondwinding aid and inner skin to form an outer skin; and cutting the outerskin to remove the second winding aids. This embodiment may also includeplacing a core piece on the inner skin.

In accordance with another embodiment of the present invention, curingfurther comprises placing a mold around an exterior of the structure;sealing the mold; placing the mold in a heating device; and applyingheat to the mold using the heating device. This embodiment may alsoinclude creating a vacuum in the mandrel and creating a vacuum in theframe mandrel.

In accordance with another embodiment of the present invention, curingfurther comprises placing a mold around an exterior of the structure;sealing the mold; placing the mold in an autoclave; and applyingpressure to the mold using the autoclave.

In accordance with an embodiment of the present invention, the structureis a fuselage of an aircraft.

In accordance with an embodiment of the present invention, a system formanufacturing a one-piece closed-shape structure using a mandrelcomprises: a preparing component configured to prepare the mandrel,wherein the mandrel comprises a bag and an armature; a first applyingcomponent configured to apply a frame mandrel to the mandrel to form aframe for the structure; a first filling component configured to fillthe mandrel and the frame mandrel with media; a second applyingcomponent configured to apply a curable resign to a fiber; a thirdapplying component configured to apply the fiber over the mandrel andframe mandrel to form the structure; a curing component configured tocure the structure; a removing component configured to remove the mediafrom the mandrel and frame mandrel; and an extracting componentconfigured to extract the mandrel and frame mandrel from the structure.

In accordance with an embodiment of the present invention, acomputer-implemented method of manufacturing a one-piece closed-shapestructure using a mandrel comprises: preparing the mandrel, wherein themandrel comprises a bag and an armature; applying a frame mandrel to themandrel to form a frame for the structure; filling the mandrel and theframe mandrel with media; applying a curable resign to a fiber; applyingthe fiber over the mandrel and frame mandrel to form the structure;curing the structure; removing the media from the mandrel and framemandrel; and extracting the mandrel and frame mandrel from thestructure.

In accordance with another embodiment of the present invention, a systemfor manufacturing a one-piece closed-shape structure using a mandrelcomprises: a preparing means for preparing the mandrel, wherein themandrel comprises a bag and an armature; an applying means for applyinga frame mandrel to the mandrel to form a frame for the structure; afilling means for filling the mandrel and the frame mandrel with media;a first applying means for applying a curable resign to a fiber; asecond applying means for applying the fiber over the mandrel and framemandrel to form the structure; a curing means for curing the structure;a removing means for removing the media from the mandrel and framemandrel; and an extracting means for extracting the mandrel and framemandrel from the structure.

In accordance with another embodiment of the present invention, aone-piece closed shape structure comprises: an outer shell formed of acomposite material; and a frame formed on an interior portion of theouter shell, the outer shell and frame being co-cured to form theone-piece closed shape structure. In this embodiment, the outer shellmay comprise an inner and outer skin. Further, in this embodiment, acore material may be located between the inner and outer skin.

In accordance with another embodiment of the present invention, aone-piece closed shape structure comprises: an outer skin formed of acomposite material; an inner skin formed of a composite material; aframe located on an interior portion of the inner skin; and a corematerial located between the inner and outer skin, wherein the outerskin, inner skin, frame, and core material have been co-cured to formthe one-piece closed shape structure.

In accordance with another embodiment of the invention, a one-pieceairplane fuselage comprises an outer skin formed of a compositematerial; an inner skin formed of a composite material; a frame locatedon an interior portion of the inner skin; and a core material locatedbetween the inner and outer skin, wherein the outer skin, inner skin,frame, and core material have been co-cured to form the one-pieceairplane fuselage. In this embodiment, the airplane fuselage may furthercomprise at least one integrally formed flange that has been co-curedwith the outer skin, inner skin, frame, and core material. In addition,this airplane fuselage may further comprise at least one integrallyformed wing attachment pocket that has been co-cured with the outerskin, inner skin, frame, core material, and flange.

Additional aspects of the invention are disclosed and defined by theappended claims. It is to be understood that both the foregoing generaldescription and the following detailed description are exemplary andexplanatory and are intended to provide further explanation of theinvention as claimed.

IV. BRIEF DESCRIPTION OF THE DRAWINGS

The accompanying drawings, which are incorporated in and constitute apart of this specification, illustrate several embodiments of theinvention and, together with the following description, serve to explainthe principles of the invention.

In the drawings:

FIG. 1A is a side view of an airplane;

FIG. 1B is a partially cut away side view of an airplane identifyingcertain features of the airplane, as shown in FIG. 1A;

FIG. 2A illustrates a conventional multi-piece composite fuselage.

As shown in FIG. 2A, conventional methods used in the industry typicallyconstruct a fuselage 200 from two or more pieces;

FIG. 2B illustrates a one-piece fuselage in accordance with anembodiment of the present invention;

FIG. 3 is a block diagram illustrating component processes formanufacturing a one-piece fuselage in accordance with an embodiment ofthe present invention;

FIG. 4 is a perspective view of a one-piece fuselage structure using thecomponent processes, as shown in FIG. 3;

FIG. 5 is a block diagram illustrating the component processes formanufacturing a one-piece integrally-stiffened fuselage in accordancewith an embodiment of the present invention;

FIG. 6 is a perspective view of a one-piece integrally-stiffenedfuselage using the components, as shown in FIG. 5;

FIG. 7 is a block diagram illustrating the components for manufacturinga one-piece integrally stiffened fuselage by a process in accordancewith an embodiment of the invention;

FIG. 8 is a block diagram illustrating alternative embodiments for theprocess of manufacturing a one-piece integrally stiffened fuselage inaccordance with the present invention;

FIG. 9A is a block diagram illustrating the processes of manufacturing aone-piece fuselage in accordance with one embodiment of the presentinvention, as shown in FIG. 8;

FIG. 9B is a flow diagram illustrating the internal mandrel flow formanufacturing a one-piece fuselage in accordance with one embodiment ofthe present invention, as shown in FIG. 9A;

FIG. 9C is a flow diagram illustrating the material flow formanufacturing a one-piece fuselage in accordance with one embodiment ofthe present invention, as shown in FIG. 9A;

FIG. 9D is a flow diagram illustrating the part flow for manufacturing aone-piece fuselage in accordance with one embodiment of the presentinvention, as shown in FIG. 9A;

FIG. 9E is a flow diagram illustrating the mold flow for manufacturing aone-piece fuselage in accordance with one embodiment of the presentinvention, as shown in FIG. 9A;

FIG. 9F is a flow diagram illustrating the core flow for manufacturing aone-piece fuselage in accordance with one embodiment of the presentinvention, as shown in FIG. 9A;

FIG. 9G is a flow diagram illustrating the process of manufacturing aone-piece fuselage in accordance with one embodiment of the presentinvention, as shown in FIG. 9A-9F;

FIG. 10A illustrates tooling preparation in accordance with anembodiment of the present invention, as shown in FIG. 9;

FIG. 10B is a cut-away view of a portion of an armature with a bag inaccordance with an embodiment of the present invention, as described inFIG. 10A;

FIG. 11A is a perspective view of an armature and bag in a form tool inaccordance with an embodiment of the present invention, as shown in FIG.10A;

FIG. 11B is a cut-away view of a portion of an armature and bag in aform tool in accordance with an embodiment of the present invention, asshown in FIG. 11A;

FIG. 12A illustrates introducing media into a mandrel in accordance withan embodiment of the present invention, as shown in FIG. 13;

FIG. 12B is a cut-away view of a portion of a mandrel filled with mediain accordance with an embodiment of the present invention, as shown inFIG. 12A;

FIG. 13 is a perspective view of installing a winding shaft in a mandrelin a form tool in accordance with another embodiment of the presentinvention, as shown in FIG. 12A-12B;

FIG. 14 illustrates a close-up perspective view of a mandrel in a formtool in accordance with an embodiment of the present invention, as shownin FIG. 13;

FIG. 15 illustrates another perspective view of a mandrel in a form toolin accordance with an embodiment of the present invention, as shown inFIG. 14;

FIG. 16A is a perspective view of the mandrel prepared for lay-up inaccordance with an embodiment of the present invention, as shown inFIGS. 12A-12B;

FIG. 16B is a cut-away view of the mandrel prepared for lay-up inaccordance with an embodiment of the present invention, as shown in FIG.16A;

FIG. 17 illustrates preparing an internal mandrel for filament windingof the inner skin in accordance with another embodiment of the presentinvention, as shown in FIG. 15;

FIG. 18 illustrates another perspective view of preparing the mandrelfor filament winding in accordance with an embodiment of the presentinvention, as shown in FIG. 17;

FIG. 19 illustrates preparing frame mandrels to be placed on a mandrelin accordance with an embodiment of the present invention, as shown inFIG. 18;

FIG. 20 illustrates preparing frame materials in accordance with anembodiment of the present invention, as shown in FIG. 9;

FIG. 21A is a perspective view of a mandrel with frame plies and framemandrels in place in accordance with an embodiment of the presentinvention, as shown in FIG. 9;

FIG. 21B illustrates frame plies on the mandrel in accordance with anembodiment of the present invention, as shown in FIG. 21A;

FIG. 21C illustrates frame plies and a frame mandrel on the mandrel inaccordance with an embodiment of the present invention;

FIG. 22 illustrates wing attachment plies being applied to a mandrel toform wing attachment pockets in accordance with an embodiment of thepresent invention, as shown in FIGS. 21A-21C;

FIG. 23 illustrates frame plies in frame recesses in a mandrel in moredetail in accordance with an embodiment of the present invention, asshown in FIGS. 21A-21C;

FIG. 24A illustrates a frame mandrel in a frame recess in a mandrel inmore detail in accordance with an embodiment of the present invention,as shown in FIGS. 21A-21C;

FIG. 24B illustrates a frame mandrel over frame plies in a frame recessin a mandrel in accordance with an embodiment of the present invention,as shown in FIGS. 21A-21C, 23, and 24A;

FIG. 25 illustrates preparing the mandrel for filament winding of theinner skin in accordance with an embodiment of the present invention, asshown in FIG. 9;

FIG. 26 illustrates applying filament to the mandrel for filamentwinding of the inner skin by a filament winding machine in accordancewith an embodiment of the present invention, as shown in FIG. 25;

FIG. 27A is a perspective view of a mandrel with a filament-wound innerskin in accordance with an embodiment of the present invention, as shownin FIGS. 21A-21C;

FIG. 27B is a cut-away view of a mandrel with a filament-wound innerskin in accordance with the embodiment of the present invention, asshown in FIG. 27A;

FIG. 28 is a side view of a mandrel with a filament-wound inner skinwith external end hoop plies in accordance with an embodiment of thepresent invention, as shown in FIG. 26;

FIG. 29 illustrates cutting a mandrel in accordance with an embodimentof the present invention, as shown in FIG. 9;

FIG. 30A is a perspective view of a mandrel with inner skin cut anddraped in accordance with an embodiment of the present invention, asshown in FIG. 27A;

FIG. 30B is a cut-away view of a mandrel with inner skin that has beencut and draped in accordance with an embodiment of the invention, asshown in FIG. 30A;

FIG. 31A illustrates machining core in accordance with an embodiment ofthe present invention, as shown in FIG. 9;

FIG. 31B is a perspective view of a mandrel with core material inaccordance with an embodiment of the present invention, as shown in FIG.9;

FIG. 31C is a cut-away view of a mandrel with core details in accordancewith an embodiment of the present invention, as shown in FIG. 31A;

FIG. 32 illustrates a portion of a mandrel with film adhesive coveringcore material in accordance with an embodiment of the present invention,as shown in FIGS. 31A-31B;

FIG. 33 illustrates preparing a mandrel for application of an outer skinby a filament winding machine in accordance with an embodiment of thepresent invention, as shown in FIG. 9;

FIG. 34 illustrates applying an outer skin to a mandrel by a filamentwinding machine in accordance with an embodiment of the presentinvention, as shown in FIG. 33;

FIG. 35A is a perspective view of a mandrel with a filament wound outerskin in accordance with an embodiment of the present invention, as shownin FIG. 9;

FIG. 35B is a cut-away view of a mandrel with a filament wound outerskin in accordance with an embodiment of the present invention, as shownin FIG. 35A;

FIG. 36A is a perspective view of a mandrel with outer skin cut anddraped in accordance with an embodiment of the present invention, asshown in FIG. 9;

FIG. 36B is a cut-away view of a mandrel with outer skin that has beencut and draped in accordance with an embodiment of the invention, asshown in FIG. 36A;

FIG. 37 illustrates the mandrel after cutting and draping of the outerskin in accordance with an embodiment of the present invention, as shownin FIG. 33;

FIG. 38A illustrates preparing a circumferential mold for a mandrel inaccordance with an embodiment of the present invention, as shown in FIG.9;

FIG. 38B is a cut-away view of a mandrel in the circumferential mold inaccordance with an embodiment of the present invention;

FIG. 39A illustrates preparing a circumferential mold with a vacuumsystem for the frame mandrels during curing in accordance with anembodiment of the present invention, as shown in FIGS. 38A-38B;

FIG. 39B illustrates a cut-away of the mandrel in the circumferentialmold with a vacuum system for the frame mandrels in accordance with anembodiment of the present invention, as shown in FIG. 39A;

FIG. 39C illustrates a vacuum port in a frame mandrel in accordance withan embodiment of the present invention, as shown in FIGS. 39A and 39B;

FIG. 39D illustrates a device for maintaining a vacuum in a framemandrel in accordance with an embodiment of the present invention, asshown in FIGS. 39B and 39C;

FIG. 40 illustrates curing a filament wound mandrel in a circumferentialmold in an oven in accordance with an embodiment of the presentinvention, as shown in FIG. 9;

FIG. 41 illustrates removing a circumferential mold from around aone-piece integrally stiffened fuselage on a mandrel in accordance withan embodiment of the present invention, as shown in FIG. 9;

FIG. 42 illustrates removing media from a mandrel in accordance with anembodiment of the present invention, as shown in FIG. 9;

FIG. 43 illustrates a one-piece integrally stiffened fuselage containedin a circumferential mold after removal of media and armature inaccordance with one embodiment of the present invention as shown in FIG.42;

FIG. 44 illustrates removing a bag from a one-piece integrally stiffenedfuselage in accordance with an embodiment of the present invention asshown in FIG. 41;

FIG. 45 illustrates a bag after removal from a mandrel in accordancewith an embodiment of the present invention, as shown in FIG. 44;

FIG. 46 illustrates removing frame mandrels from a one-piece integrallystiffened fuselage in accordance with an embodiment of the presentinvention as shown in FIG. 9;

FIG. 47 illustrates a one-piece integrally-stiffened fuselagemanufactured in accordance with one embodiment of the present inventionas shown in FIG. 9;

FIG. 48 illustrates a one-piece integrally stiffened fuselagemanufactured in accordance with one embodiment of the present inventionas shown in FIG. 9;

FIG. 49 is a block diagram illustrating the process of manufacturing aone-piece fuselage in accordance with another embodiment of the presentinvention, as shown in FIG. 9;

FIG. 50 illustrates assembling a circumferential mold around a mandrelin accordance with an embodiment of the present invention, as shown inFIG. 49;

FIG. 51 illustrates bagging a circumferential mold in accordance with anembodiment of the present invention, as shown in FIG. 50;

FIG. 52 illustrates placing a circumferential mold in an autoclave forcuring in accordance with an embodiment of the present invention, asshown in FIG. 51;

FIG. 53 illustrates removing a circumferential mold after curing in anautoclave in accordance with an embodiment of the present invention, asshown in FIG. 52; and

FIG. 54 illustrates a one-piece integrally-stiffened fuselagemanufactured in accordance with another embodiment of the presentinvention, as shown in FIG. 49.

V. DESCRIPTION OF THE EMBODIMENTS

A. Introduction

Methods and structures in accordance with the present invention will nowbe described with respect to an embodiment of a one-piece structure, anaircraft fuselage. The attached figures illustrate the manufacture ofboth a fuselage containing a tail cone and a fuselage without a tailcone. The invention as claimed, however, is broader than fuselages andextends to other closed-shape structures, such as, other aircraft,automotive, forklift, or watercraft structures.

B. Methods and Structures

FIG. 1A is a side view of an airplane. As shown in FIG. 1A, airplane 100consists of engine section 101, fuselage 102, empennage 103, and wings104. Engine section 101, empennage 103, and wings 104 connect tofuselage 102. Airplane 100 may be any type of airplane, such as, prop,jet, or other type. Airplane 100 is the type of airplane for which aone-piece fuselage could be constructed (which is described in moredetail below). In one implementation, the one-piece fuselage includesfuselage 102. In other implementations, the one-piece fuselage may alsoinclude engine section 101, empennage 103, wings 104, and/or any otherparts of aircraft 100 (not shown). This implementation is merelyexemplary, and other implementations may also be used.

FIG. 1B is a partially cut away side view of an airplane identifyingcertain features of the airplane, as shown in FIG. 1A. In FIG. 1B,airplane 100 is described in more detail than in FIG. 1A. As shown inFIG. 1B, engine section 101 contains an engine 107, an engine mount 108,and a firewall 109. Engine section 101, engine mount 108, and firewall109 are all connected to fuselage 102. In some implementations, engine107 is connected to fuselage 102 via engine mount 108. However, otherimplementations may have engine 107 connected directly to fuselage 102.Further, engine mount 108 may be a separate component, as shown in FIG.1B, or engine mount 108 may be a part of either engine section 101 orfuselage 102. These implementations are merely exemplary, and otherimplementations may also be used.

Empennage 103 contains tail cone 106, vertical stabilizers 107, andhorizontal stabilizers 108. Empennage 103 may be a separate component ofairplane 100, as shown in FIG. 1B, or tail cone 106 may be a part offuselage 102 with vertical stabilizer 107 and horizontal stabilizer 108being separate pieces. These implementations are merely exemplary, andother implementations may also be used.

Wings 104 generally include left wing 121 (shown) and right wing 122(not shown). Wings 104 are connected to fuselage 102 by wing spars 110.Wing spars 110 support wings 101 within fuselage 102. Other wingconfigurations may be used for airplane 100, such as a bi-wingconfiguration or a tri-wing configuration or other wing configurations.In addition, a canard (not shown) and winglets (not shown) may also beused with airplane 100. Also, airplane 100 depicts a low wing aircraft,but airplane 100 may also be a high-wing, mid-wing, or other wing-designaircraft.

Fuselage 102 contains panel section 116, seat supports (not shown),seats 112, access doors 115, luggage access doors 113, and windows 114.Panel section 116 holds flight instruments for airplane 100. Seatsupports hold seats 112. Access door 115 is depicted as a single door onthe left side of airplane 100, as shown in FIG. 1B. Access door 115 mayalso be located on the right side of airplane 100. Further, additionalor other doors may be included within a group of access doors 115, suchas a second set of access doors or any other access door configurations.Luggage access door 113 is depicted as located on the back left side ofairplane 100, as shown in FIG. 1B, but luggage access door 113 may belocated anywhere on airplane 100. In addition, airplane 100 may containmultiple luggage access doors 113. Fuselage 102 also contains wing sparattachment boxes (not shown).

Windows 114 include front window 117 and side windows 118. FIG. 1Bdepicts two side windows 118, but other configurations may be used forside windows 118 such as one side window 118 or two or more side windows118. Windows 114 may also include a rear window (not shown). Windows 114may also include other windows, such as skylight windows (not shown),camera windows (not shown), or any other type of window.

As shown in FIG. 1B, fuselage 102 has numerous openings, such as accessdoors 115, windows 114, and luggage access doors 113. Also, there areother openings that are not shown such as an engine mount block (notshown) for engine mount 108, an empennage mounting block (not shown) forempennage 106, and landing gears mounts (not shown) for landing gear119.

Because of these openings, portions of fuselage 102 may be strengthenedfor support around these openings. For example, window frames 140 may bestrengthened to support windows 114. Other areas may also bestrengthened, such as seat supports 130 (not shown, but describedabove). Other strengthening may also be necessary for the engine mountblock (not shown), the landing gear mounts (not shown), the empennagemounting block (not shown), and roll-over frames (not shown). Stillother areas may also need to be strengthened, depending on the design ofairplane 100. These implementations are merely exemplary, and otherimplementations may also be used.

FIG. 2A illustrates a conventional multi-piece composite fuselage. Asshown in FIG. 2A, conventional methods used in the industry typicallyconstruct a fuselage 200 from two or more pieces. Those pieces consistof fuselage halves 205 and 210, frame stiffening structures for thepassenger area 220, and wing spar attachment boxes 230. In compositeaircraft manufacture, these pieces are typically manufactured fromfiberglass prepreg. These conventional methods also require the steps ofbonding the pieces together, machining of the pieces at the joint areas,machining the core frames, and various other mechanical assemblyprocesses.

FIG. 2B illustrates a one-piece fuselage in accordance with anembodiment of the present invention. As shown in FIG. 2B, a fuselage 250is a one-piece structure, including exterior surface of the fuselage260, frame sections 270, attachment pockets for the wings 280, and otherframes sections, attachments pockets, and flanges (not shown). Notably,the use of a one-piece fuselage eliminates the assembly operations thatare associated with the conventional methods for manufacturing afuselage as well as providing other advantages, as described above. Thisimplementation is merely exemplary, and other implementations may alsobe used.

FIG. 3 is a block diagram illustrating component processes formanufacturing a one-piece fuselage in accordance with an embodiment ofthe present invention. As shown in FIG. 3, the component processes formanufacturing a one-piece fuselage 300 include molding 310, tooling 320,tooling integration 330, and fiber placement 340. Molding 310 includesthe use of any type of molding. For example, molding 310 may includesuch things as hand lay up of fiberglass or graphite prepreg into molds,pressing of sheet molding compounds, injection of molding compounds intodies, and/or machine lamination of composite prepreg onto molds.

Tooling 320 includes the use of any type of tooling needed for molding.For example, tooling 320 may include the use of metal molds, molds madefrom composite materials, and/or mandrels made from metals and compositematerials. Tooling 320 also includes tooling made from elastomericmaterials such as silicone, urethane, or natural rubbers. Tooling 320further includes the use of plastic or metal dies and punches.

Tooling integration 330 includes any combination of molding 310 withtooling 320. For example, tooling integration 330 includes vacuumsealing of a part cavity, pressurization of tool cavities, and/orapplication of vacuum pressure in tool cavities.

Fiber placement 340 includes any placement using any form of fiber. Forexample, fiber placement 340 includes such things as winding with carbontape, winding with carbon tow, winding with glass fiber or roving,winding with glass tape, wrapping of glass or carbon prepreg materials,and/or wrapping of carbon and glass fiber materials. As shown in FIG. 3,molding 310, tooling 320, tooling integration 330, and fiber placement340 may be used to create one-piece fuselage 300. This implementation ismerely exemplary, and other implementations may also be used.

FIG. 4 is a perspective view of a one-piece fuselage structure using thecomponent processes, as shown in FIG. 3. As shown in FIG. 4, a one-piecefuselage 400 may be created using the processes described in FIG. 3. Forexample, molding 310 is used to create such things as molded frames 410,molded integral flanges 415 (for attachment of the bulkhead), and moldedintegral wing attachment hard points 420 (for attachment of the wings).Tooling 320 is used to create such things as armature 425 and mandrel430 (placed inside the fuselage). Tooling integration 330 is used tointegrate tooling 320 with molding 310 to prepare for fiber placement340 of fuselage 400. Fiber placement 340 creates fuselage skin 450. FIG.4 depicts just some examples of the uses of the components of FIG. 3 ina one-piece fuselage, and many other uses may be made of thesecomponents just some of which are described herein with reference toFIG. 4).

FIG. 5 is a block diagram illustrating the component processes formanufacturing a one-piece integrally-stiffened fuselage in accordancewith an embodiment of the present invention. As shown in FIG. 5, aone-piece integrally-stiffened fuselage 500 includes molding co-curedhollow and foam-filled frames 510, elastomeric tooling for internalfuselage mandrel 520, tooling integration for integrally-stiffenedfuselage 530, and fiber placement fuselage shape 540.

For fuselage 500, molding co-cured hollow and foam filled frames 510includes the molding of stiffening structure inside of the fuselageshell that is co-cured with that shell. Such molded structure may alsoinclude flanges that are integral with the shell. Other moldedstructures may further include wing attachment pockets and pockets forengine truss mount fittings.

Elastomeric tooling for internal fuselage mandrel 520 includes the useof an elastomeric tooling associated with molding the internal shape ofthe fuselage. In this context, elastomeric tooling refers to a mandrelthat is used to maintain the internal shape of the fuselage during framelay up and filament winding.

Tooling integration for integrally-stiffened fuselage 530 involves thejoining of molding 510 and elastomeric tooling 520 in a manner thatproduces the fuselage shape. Tooling integration 530 includes suchthings as application of a vacuum and/or pressure in various moldcavities to obtain the desired fuselage shape.

Finally, fiber placement fuselage shape 540 is used to create thefuselage skin material and shape. During fiber placement 540, fiber iswound directly on the elastomeric mandrel to create the fuselage skin.Fiber placement 540 also includes the creation of skin material bywinding fiber over a secondary mandrel and then cutting ply pieces toobtain frame and integral stiffening structures.

Co-curing of fuselage components creates integral stiffening forone-piece integrally stiffened fuselage 500. The co-curing of thesecomponents reduces the chances of defects created during manufacturingof separate elements and their subsequent joining using bonding ormechanical fasteners. The co-cured integral stiffening distributes loadsmore uniformly throughout the fuselage during flight and in the event ofan off field landing.

FIG. 6 is a perspective view of a one-piece integrally-stiffenedfuselage using the components, as shown in FIG. 5. As shown in FIG. 6, aone-piece fuselage 600 may be created using the components described inFIG. 5. For example molding 510 may be used to create co-cured hollowand foam filled frames 610, integral flanges 615 (such as for bulkheadattachment), and integral wing attachment hard points 620.

Tooling 520 may be used to create armature 625 and a reusableelastomeric mandrel 630 both of which go inside of fuselage 600. Tooling520 may also include filament-wound broad goods. In general, broad goodsare custom-sized pieces of composite materials, and filament-wound broadgoods are custom-sized pieces of composite materials that have beenfilament-wound. For example, a filament-wound broad good would become aco-cured hollow frame ply. These filament-wound broad goods may be usedfor integral frames 610, flanges 615, longerons 610, and doublers 620.

Tooling integration 530 may be used to integrate such things as toolingfor fuselage 600 and tooling for the filament winding of fuselage skins(such as fuselage skin 650) with molding 510.

Fiber placement 540 may be used to create fuselage skin plies 650. FIG.6 depicts just some examples of the uses of the components of FIG. 5 ina one-piece integrally-stiffened fuselage, and many other uses may bemade of these components Oust some of which are described with referenceto FIG. 6).

FIG. 7 is a block diagram illustrating the components for manufacturinga one-piece integrally stiffened fuselage by a process in accordancewith an embodiment of the invention. As shown in FIG. 7, process forcreation of one-piece integrally-stiffened fuselage 700 includes moldingco-cured hollow and foam filled frames 510, elastomeric tooling forinternal fuselage mandrel 520, tooling integration 530, and fiberplacement fuselage shape 540. As shown in FIG. 7, a process for thecombination of molding co-cured hollow and foam filled frames 510,elastomeric tooling for internal fuselage mandrel 520, toolingintegration 530, and filament placement fuselage shape 540 may result inone-piece integrally-stiffened fuselage 500. This implementation ismerely exemplary, and other implementations may also be used.

FIG. 8 is a block diagram illustrating alternative embodiments for theprocess of manufacturing a one-piece integrally stiffened fuselage inaccordance with the present invention. As shown in FIG. 8, process forcreation of one-piece integrally-stiffened fuselage (as described inFIG. 7) includes Process Alternate #1 805, Process Alternate #2 810, andOther Process Alternates 815. Other Process Alternates 815 shows thatvarious alternative processes may be used for creating one-pieceintegrally-stiffened fuselage 700. Process Alternate #1 805 is depictedin FIGS. 9-48. Process Alternate #2 810 is depicted in FIGS. 49-53.Although not depicted in separate figures, Other Process Alternates 815show that processes other than those described herein may be used forthe process for creation of one-piece integrally-stiffened fuselage 800.

1. Alternate 1

FIG. 9A is a block diagram illustrating the processes of manufacturing aone-piece fuselage in accordance with one embodiment of the presentinvention, as shown in FIG. 8. As shown in FIG. 9A, process for creationof one-piece integrally stiffened fuselage 900 includes five processesincluding internal mandrel flow 901, material flow 902, part flow 903,mold flow 904, and core flow 905. Internal mandrel flow 901 is depictedin FIG. 9B. Material flow 902 is depicted in FIG. 9C. Part flow 903 isdepicted in FIG. 9D. Mold flow 904 is depicted in FIG. 9E. Core flow 905is depicted in FIG. 9F. FIG. 9G illustrates the integration of flows901-905.

FIG. 9B is a flow diagram illustrating the internal mandrel flow formanufacturing a one-piece fuselage in accordance with one embodiment ofthe present invention, as shown in FIG. 9A. As shown in FIG. 9B,internal mandrel flow 901 describes the flow of the internal tooling forcreation of the fuselage. Internal mandrel flow 901 begins with toolingpreparation 907. Tooling preparation 907 includes placing of an armatureinside of a bag, forming the bag inside of a form tool, and filling thespace between the bag and the form tool with media to form the internalmandrel. The mandrel exterior (which is described in detail on thefollowing figures) is in the shape of the desired fuselage interior.

After tooling preparation 907, the following actions take place: placeframes and frame mandrels 908, wind inner skin 910, cut and drape 911,place core 912, wind outer skin 913, cut and drape 914, close mold 915,cure 916, demold 917 and extract frame mandrels 918. Each of theseactions are described in detail below (see FIGS. 10-48).

FIG. 9C is a flow diagram illustrating the material flow formanufacturing a one-piece fuselage in accordance with one embodiment ofthe present invention, as shown in FIG. 9A. As shown in FIG. 9C,material flow 902 describes the flow of fiber and resin for creating thecomposite material that becomes one-piece integrally-stiffened fuselage500. Material flow 902 begins with inspect incoming materials 930.Inspect incoming materials 930 involves inspection of the fiber andresin used by system 900. Fiber and resin at inspect incoming materials930 are inspected for conformity for use in prepare frame materials 932and filament winder 935.

Prepare frame materials 932 involves the preparation of a fiber andresin, which includes filament winding of broad goods, cutting with aply cutter, and the preparation of frame and doubler plies. Filamentwinder 935 includes loading of the fiber and resin into a filamentwinder device, such as those known in the art. Filament winder 935creates an inner skin in wind inner skin 910. Filament winder 935 alsocreates an outer skin in wind outer skin 913. Each of these actions isdescribed in detail below (see FIGS. 10-48).

In material flow 902, any type of fiber and any type of resin may beused. Some of the fibers that are found to be acceptable include: TorayT700, T600, and T300 that are available in 3 K, 6 K, and 12 K tow count,Amoco T-300 and T-650 that are available in 3 K, 6 K, and 12 K towcount, Hexcal AS4 that is available in 3 K and 6 K tow count, Fortafil3(C), Grafil 34-600WD, and Panex 33. In one implementation, fibers thatare never twisted may be used, although other fibers may also be used inother implementations. In most implementations, an acceptable tow countfor the spool is dependent upon part size.

Additionally, any type of curable resin may be used. Some curable resinsthat have been found acceptable include epoxy resin with a roomtemperature viscosity of 10,000 to 125,000 cps. The viscosity to be usedgenerally depends upon the shape of the part being filament wound. Anytype of Shell epoxy resin may also be used. Shell epoxy resins that havebeen found acceptable include combinations of 862 and 1050 with “W”curative and accelerator 537. In addition, Shell epoxy resins that havebeen found acceptable may use tougheners from Nippon zeon, which havebeen shown to have desirable physical properties. Shell epoxy resins canbe used separately or in combination with other resins to obtain thedesired properties. Furthermore, still other resins may also be used.For example, high tack resins may be used under certain circumstances(such as holding fibers in place during winding). Therefore, any type ofcurable resin and any combination of curable resin types may be used.

FIG. 9D is a flow diagram illustrating the part flow for manufacturing aone-piece fuselage in accordance with one embodiment of the presentinvention, as shown in FIG. 9A. As shown in FIG. 9D, part flow 903describes the flow of the part (i.e. the fuselage) during themanufacturing process. Part flow 903 includes most of the actions ofinternal mandrel flow 901. Part flow 903 begins with prepare framematerials 932. Part flow 903 then includes the following actions: placeframes and frame mandrels 908, wind inner skin 910, cut and drape 911,place core 912, wind outer skin 913, cut and drape 914, close mold 915,cure 916, de-mold 917, extract frame mandrels 918, visually inspect part919, trim 920, prime and paint 921, and store for assembly 922. Partflow 903 also includes monitoring process 936, monitoring process 937,and monitoring process 978 for monitoring wind inner skin 910, windouter skin 913, and cure 916, respectively. Finally, part flow 903 alsoincludes prepare oven 975, and cut-outs to quality control 985. Each ofthese actions is described below (see FIGS. 10-48).

FIG. 9E is a flow diagram illustrating the mold flow for manufacturing aone-piece fuselage in accordance with one embodiment of the presentinvention, as shown in FIG. 9A. As shown in FIG. 9E, mold flow 904describes the actions involving a mold during the manufacturing process.In one implementation, a mold (which will be described in more detailbelow) is in the shape of a circumferential mold for holding (thereforemolding) the structure inside the circumferential mold. In oneimplementation, the circumferential mold is made of metal. In otherimplementations, the circumferential mold may be made out of any othermaterials. Mold flow 904 begins with prepare mold 940. After preparemold 940, close mold 915 occurs. Close mold 915 includes closing themold. In some implementations, close mold 915 may also include applyinga vacuum to the mold or applying pressure to the mold. After close mold915, cure 916 occurs (which in some implementations may be preceded byprepare oven 975). In one implementation, cure 916 involves heating themold, or in another implementation cure 916 involves pressurization ofthe mold. In other implementations, cure 916 may include curing by anyother method. After cure 916, demold 917 occurs. Demold 917 includes theremoval of the contents of the mold. After demold 917, the mold isreused in prepare mold 940. All of these actions will be described below(see FIGS. 10-48).

FIG. 9F is a flow diagram illustrating the core flow for manufacturing aone-piece fuselage in accordance with one embodiment of the presentinvention, as shown in FIG. 9A. As shown in FIG. 9F, core flow 905involves the core materials that give buckling stiffness to the desiredstructure, such as one-piece integrally-stiffened fuselage 500. Coreflow 905 begins with inspect incoming materials 930. At inspect incomingmaterials 930, core material (such as honeycomb core material) isinspected for use. After inspect incoming materials 930, the corematerial is machined to shape in machine core to shape 952. Aftermachine core to shape 952, the core material is cleaned in clean core950. After clean core 950, the core material is used by system 900 (SeeFIG. 9G) in place core 912. Notably, place core 912 may involvemanipulation of the shape of the desired structure by cut and drape 911.For example, between cut and drape 911 and place core 912, additionalmaterial may be added to the core material for manipulation of the shapeof the desired structure. Any type of core material may be used for thecore material. These actions will be described below (see FIGS. 10-48).

FIG. 9G is a flow diagram illustrating the process of manufacturing aone-piece fuselage in accordance with one embodiment of the presentinvention, as shown in FIG. 9A-9F. FIG. 9G illustrates the combinationin system 900 of internal mandrel flow 901, material flow 902, part flow903, mold flow 904, and core flow 905, described in FIGS. 9A-9G.Internal mandrel flow 901 is depicted in FIG. 9G as a line with arrows.Material flow 902 is depicted in FIG. 9G as a line with arrows withcircles through the line. Part flow 903 is depicted in FIG. 9G as a boldline with arrows. Mold flow 904 is depicted in FIG. 9G as a line witharrows with circles next to the line. Core flow 905 is depicted in FIG.9G as a line with arrows with slash marks through the line.

As referenced above, the actions taken in system 900 are described inFIGS. 10-48. FIGS. 10-48 illustrate the use of system 900 to create botha one-piece integrally stiffened fuselage with a tail cone and anintegrally stiffened fuselage without a tail cone. Therefore, FIGS.10-48 illustrate the components of system 900 being used to createmultiple structures. Those figures illustrating the creation of aone-piece integrally stiffened fuselage with a tail cone will benumbered consistently with one another and those figures illustratingthe creation of a one-piece integrally stiffened fuselage without a tailcone will be number consistently with one another.

As shown in FIG. 9G, system 900 begins with tooling preparation 907.Tooling preparation 907 is described in FIG. 10A.

FIG. 10A illustrates tooling preparation in accordance with anembodiment of the present invention, as shown in FIG. 9. FIG. 10Aprovides an example of tooling preparation 907 from FIG. 9. FIG. 10Ashows the preparation of tooling, such as a mandrel 1000. In oneimplementation, mandrel 1000 is a reusable elastomeric mandrel, such asthat currently available through International Design Technologies, Inc(IDT). However, any mandrel may be used.

Mandrel 1000 may include a bag 1010 and an armature 1020. Bag 1010 maycomprise premolded silicone, or bag 1010 may consist of any other formor substance. Some silicone bag materials that have been foundacceptable include those available from Arlon-Mosite, Kirkhill, andD-Aircraft Products SMC 950. In addition, there are many other suppliersof high temperature (up to 400° F.), unfilled, and uncured siliconesheet materials that may be used, depending upon the cure temperature ofthe desired part.

Armature 1020 may be made of any material. In one implementation, awelded metal armature is used. However, other materials could be used toform the armature. To minimize weight and mandrel bending, armature 1020may be as large as possible, while allowing it to be removed from bag1010 and from the completed fuselage. These implementations are merelyexemplary, and other implementations may also be used.

FIG. 10B is a cut-away view of a portion of an armature with a bag inaccordance with an embodiment of the present invention, as described inFIG. 10A. As shown in FIG. 10B, armature 1020 is placed through bag 1010to form cavity 1030. The space difference between armature 1020 and bag1010 provides for cavity 1030. To form cavity 1030, bag 1010 is sealedat each end to armature 1020. In one implementation, clamps and/or boltsare used to seal each end of bag 1010. Armature 1020 thus supports bag1010. Notably, bag 1010 has a desired pre-molded shape 1040. Bag 1010may lack the rigidity to maintain desired shape 1040 without supportfrom additional tooling. Therefore, as described below, additionaltooling may be used to maintain desired shape 1040. Theseimplementations are merely exemplary, and other implementations may alsobe used.

FIG. 11A is a perspective view of an armature and bag in a form tool inaccordance with an embodiment of the present invention, as shown in FIG.10A. As shown in FIG. 11A, following placement of armature 1020 in bag1010 (as described in FIG. 10A), armature 1020 and bag 1010 are placedin a form tool 1110 and bag 1010 is sealed at both ends to form tool1110. In one implementation, form tool 1110 covers most of armature 1020and bag 1010. Form tool 1110 provides a desired shape to outside surfaceof bag 1010. This implementation is merely exemplary, and otherimplementations may also be used.

FIG. 11B is a cut-away view of a portion of an armature and bag in aform tool in accordance with an embodiment of the present invention, asshown in FIG. 1A. As shown in FIG. 11B, bag 1010 is between form tool1110 and armature 1020. Bag 1010 is sealed at each end to both form tool1110 and armature 1020 to form enclosed cavity 1030 and enclosed cavity1120. Enclosed cavity 1030 is between outside surface of armature 1020and inside surface of bag 1010 and enclosed cavity 1120 is betweeninside surface of form tool 1110 and outside surface of bag 1010. In oneimplementation, form tool 1110 is equipped with ports for enclosedcavity 1120 (not shown) to control pressure and quantity of air withinenclosed cavity 1120. Enclosed cavity 1030 may also be equipped withports (not shown) to control pressure and quantity of air withinenclosed cavity 1030. These implementations are merely exemplary, andother implementations may also be used.

In one implementation, to provide the desired shape to outside surfaceof bag 1010, the air is vented from enclosed cavity 1120 through ports1130 (not shown) while pressurized air is inserted into enclosed cavity1030 through ports (not shown) forcing outside surface of bag 1010against inside surface of form tool 1110. Ports for enclosed cavity 1120are then sealed to maintain outside surface of bag 1010 against insidesurface of form tool 1110. Ports to enclosed cavity 1030 may then bekept pressurized or they may be vented to the atmosphere. In anotherimplementation, to provide desired shape to outside surface of bag 1010,the air is evacuated from enclosed cavity 1120 through ports while portsinto enclosed cavity 1030 are left vented to the atmosphere which forcesoutside surface of bag 1010 against inside surface of form tool 1110.These implementations are merely exemplary, and other implementationsmay also be used.

FIG. 12A illustrates introducing media into a mandrel in accordance withan embodiment of the present invention, as shown in FIG. 13. As shown inFIG. 12A, after forming bag 1010 to the shape of form tool 1110 (asdescribed in FIGS. 11A-11B), media 1210 may be introduced into enclosedcavity 1030 to provide the desired shape to bag 1010 (not shown, butshown in FIG. 12B). In one implementation, media 1210 is introducedthrough a sealable opening (not shown) inside armature 1020. Media 1210may be any material used to provide rigidity to bag 1010. In oneimplementation, media 1210 is a lightweight insulator material, such asporous ceramic materials used for water filtration. In anotherimplementation, aluminum hollow-beaded materials may be used. Theseimplementations are merely exemplary, and other implementations may alsobe used.

In one implementation, when media 1210 is introduced into enclosedcavity 1030, it may be introduced under pressure if enclosed cavity 1030is pressurized, under atmospheric conditions if enclosed cavity 1030 isvented to atmosphere, or under less than atmospheric conditions ifenclosed cavity 1030 is maintained under some pressure less thanatmospheric. As shown in FIG. 12A, in one implementation, theintroduction of media 1210 is performed in a semi-horizontalorientation. However, in other implementations, other orientations, suchas a vertical orientation or any other orientation, may be used forintroducing media 1210 into enclosed cavity 1030. These implementationsare merely exemplary, and other implementations may also be used.

FIG. 12B is a cut-away view of a portion of a mandrel filled with mediain accordance with an embodiment of the present invention, as shown inFIG. 12A. As shown in FIG. 12B, media 1210 is introduced into enclosedcavity 1030, which is between armature 1020 and bag 1010, as heldtogether by form tool 1110. After the introduction of media 1210, media1210 may be compacted to settle the media. In one implementation, thecompacting of media 1210 occurs by vibrating form tool 1110. In anotherimplementation, compacting of media 1210 occurs by tamping media 1210.These implementations are merely exemplary, and other implementationsmay also be used.

Following compaction of media 1210 the air within enclosed cavity 1030may be removed as completely as possible to complete a pressuredifference between enclosed cavity 1030 and the atmosphere. Thispressure difference causes bag 1010 to retain its shape once form tool1110 is removed. If a pressure difference between enclosed cavity 1030and the atmosphere is not maintained, bag 1010 may lose the desiredshape established by form tool 1110. In one implementation, five poundsper square inch (psi) of pressure difference between enclosed cavity1030 and atmospheric pressure has been demonstrated sufficient to causebag 1010 to retain the desired shape. This implementation is merelyexemplary, and other implementations may also be used.

FIG. 13 is a perspective view of installing a winding shaft in a mandrelin a form tool in accordance with another embodiment of the presentinvention, as shown in FIGS. 12A-12B. FIG. 13 is also the first drawingillustrating the manufacture of a mandrel without a tail cone. As shownin FIG. 13, following the introduction of media into mandrel 1390, awinding shaft 1330 is inserted into mandrel 1300. In one implementation,mandrel 1390 incorporates armature 1320, winding shaft 1330, compactedmedia (not shown), and bag (not shown). Winding shaft 1330 is used torotate mandrel 1390 during fiber placement. In one implementation,winding shaft 1330 is inserted into a box channel within armature 1320.

As shown in FIG. 13, mandrel 1390 is surrounded by form tool 1310, asdescribed above. Form tool 1310 incorporates pivot 1340. In oneimplementation, pivot 1340 allows form tool 1310 and mandrel 1300 torotate to vertical, if needed. Form tool 1310 also includes clamps 1350,bolts 1360, vacuum port 1370, and end plates 1380. In oneimplementation, clamps 1350 are used to seal a bag (not shown) aroundform tool 1310. In this implementation, bolts 1360 are used to join andseal segments of form tool 1310 to each other. In addition, bolts 1360may also be used to seal end plates 1380 to bag 1420, form tool 1310 andto armature 1320.

FIG. 14 illustrates a close-up perspective view of a mandrel in a formtool in accordance with an embodiment of the present invention, as shownin FIG. 13. As shown in FIG. 14, one section of form tool 1310 has beenremoved from around mandrel 1390. As shown in FIG. 14, the externalsurface of bag 1420 is formed to the shape of internal surface of formtool 1310.

In one implementation, form tool 1310 includes a vacuum port 1430.Vacuum port 1430 connects to an interior surface of form tool 1310.Vacuum port 1430 is used to vent or remove air from between interiorsurface of form tool 1310 and exterior surface of bag 1420.

FIG. 15 illustrates another perspective view of a mandrel in a form toolin accordance with an embodiment of the present invention, as shown inFIG. 14. As shown in FIG. 15, a section of form tool 1310 has beenremoved from mandrel 1390. As further shown in FIG. 15, bag 1420 retainsthe desired shape imparted to it by form tool 1310. In oneimplementation, a pressure differential is maintained between amedia-filled enclosed cavity situated between armature 1320 and bag 1420and the atmosphere. This implementation is merely exemplary, and otherimplementations may also be used.

As shown in FIG. 15, mandrel 1390 includes frame recesses 1540 and wingattachment pocket recesses 1550. Frame recesses 1540 and wing attachmentpocket recesses 1550 are located on the external surface of bag 1420. Inone implementation, frame recesses 1540 and wing attachment pocketrecesses 1550 are created by the inside surface of form tool 1310.However, it may be problematic for form tool 1310 to create framerecesses 1540 and wing attachment pocket recesses 1550 because of thetendency of form tool 1310 to have either no draft or negative draft.For this reason, the removal of form tool 1310 could be difficult fromaround certain portions of mandrel 1390.

In one implementation, these problems are overcome by making form tool1310, as shown in FIG. 15, in multiple pieces having required draft. Inanother implementation, the negative draft features are made as separatedetails that fit within recesses in the inside of the form tool and aredetachable from outside of the form tool when it is necessary to removeform tool from around formed mandrel. These implementations are merelyexemplary, and other implementations may also be used.

FIG. 16A is a perspective view of the mandrel prepared for lay-up inaccordance with an embodiment of the present invention, as shown inFIGS. 12A-12B. As shown in FIG. 16A, following introduction of internalmedia 1210 (described in FIGS. 12A-12B), form tool 1110 is removed toexpose mandrel 1610. Mandrel 1610 is then cleaned and prepared forlay-up. Lay-up is the procedure of applying composite materials atdesired locations to the exterior surface of the formed mandrel. Thesematerials may (when cured) form stiffening structure, frames, within thefuselage, or when placed following placement of an inner skin, as coredetails, add buckling strength to the fuselage skin.

In one implementation, mandrel 1610 contains frame recesses 1620, windowrecesses 1630, door recesses 1640, core detail recesses 1650, and wingattachment pocket recesses 1660. These recesses are used to formfeatures such as frames, windows, doors, core pockets, and wingattachment pockets in the fuselage. This implementation is merelyexemplary, and other recesses and other implementations may also beused.

FIG. 16B is a cut-away view of the mandrel prepared for lay-up inaccordance with an embodiment of the present invention, as shown in FIG.16A. As shown in FIG. 16B, mandrel 1610 includes armature 1020, bag1010, and enclosed cavity 1030 filled with media 1210. As shown in FIG.16B, the insertion of media 1210 into enclosed cavity 1030 andsubsequent evacuation of air has caused the outside surface of bag 1010to hold the desired inside surface shape of the removed form tool 1110.This “shape memory” provides for features desired for lay-up of windowdoublers, door frames, core details, and wing attachment lugs. Thisimplementation is merely exemplary, and other implementations may alsobe used.

FIG. 17 illustrates preparing an internal mandrel for filament windingof the inner skin in accordance with another embodiment of the presentinvention, as shown in FIG. 15. As shown in FIG. 17, mandrel 1390 isplaced in a winding cart 1710 in preparation for filament winding.Winding end aids 1720 are positioned on the ends of mandrel 1500.Winding end aids 1720 are used to eliminate concave or flat windingsurfaces. Other winding aids are also depicted in FIG. 17. Examples ofother winding aids include frames 1740, door recess fillers 1760, andwindshield area fillers 1750. Other winding aids provide a surface uponwhich the filament winding machine places the fibers (and resin) so thatthe fibers do not shift as the mandrel is rotated. Other winding aids1725 are also used to protect mandrel 1390 from being cut during cut anddrape 911. Other winding aids 1725 may further include guide features toguide cutting of plies during cut and drape 911 (described below). Inaddition, gap winding aids 1727 (not shown) are also used in areas whereother winding aids 1725 are higher than mandrel 1390. These gap windingaids 1727 ensure that all surfaces are convex prior to filament winding.

FIG. 18 illustrates another perspective view of preparing the mandrelfor filament winding in accordance with an embodiment of the presentinvention, as shown in FIG. 17. As shown in FIG. 18, gap winding aids1727 have been installed to make all surfaces convex. Gap winding aids1727 are used because the excessive concave area on mandrel 1390 wouldmake filament winding difficult.

FIG. 19 illustrates preparing frame mandrels to be placed on a mandrelin accordance with an embodiment of the present invention, as shown inFIG. 18. As shown in FIG. 19, frame mandrel tools 1910 are used tocreate desired shapes using frame mandrels 1920. In one implementation,frame mandrels 1920 are held to the inside shape frame mandrel tools1910 by drawing a vacuum between the inside surface of form tools 1910and the outside surface of frame mandrels 1920. Frame mandrels 2020 arethen filled with media. Finally, this media is compacted to give theframe mandrel the desired shape.

In one implementation, to maintain the desired shape, the air within theframe mandrel cavity, which has been completely filled with media, isevacuated. This causes media to lock together retaining the form toolshape. In another implementation, frame mandrels may include armatures.These implementations are merely exemplary, and other implementationsmay also be used.

FIGS. 10-19 have described tooling preparation 907, as shown in FIG. 9.As shown in FIG. 9, following tooling preparation 907, prepare framematerials 932 occurs. Prepare frame materials 932 is described in FIG.20.

FIG. 20 illustrates preparing frame materials in accordance with anembodiment of the present invention, as shown in FIG. 9. As shown inFIG. 20, frame material 2000 may be cut to produce ply pieces which willbe formed and placed on a mandrel in a predetermined location to produceframes. Frame material 2000 may also be cut to produce ply pieces, whichwill be formed and placed on a mandrel in predetermined locations toproduce integral doublers, longerons, flanges, and attachment lugs.

Frame material 2000 may be used for frame plies 2010, doubler plies2020, longeron plies 2030, integral flange plies 2040, and wingattachment pocket plies 2050. Frames, doublers, longerons, flanges, andattachment pockets are structures that enhance the strength and utilityof the fuselage. Frame material 2000 may include prepreg fabric orfilament-wound broad goods. Other frame material may also be used.

FIG. 20 has described prepare frame materials 932, as shown in FIG. 9.As shown in FIG. 9, following prepare frame materials 932, place framesand frame mandrels 908 occur. Place frames and frame mandrels 908 aredescribed in FIGS. 21A-24B.

FIG. 21A is a perspective view of a mandrel with frame plies and framemandrels in place in accordance with an embodiment of the presentinvention, as shown in FIG. 9. As shown in FIG. 21A, mandrel 1610includes recesses for plies, including frame plies 2010, integral flangeplies 2040, and wing attachment pocket plies 2050. In oneimplementation, these frame plies 2010 are placed in frame recesses 1620(see FIG. 21B) on mandrel 1610. Once frame plies 2010 have been placed,frame mandrels 1920 may be placed. In one implementation, frame mandrels1920 (as described in FIG. 19) may be placed upon frame plies 2010 orflange plies 2040 to provide the support during the cure process. In oneimplementation, mandrel 1610 may be placed in winding cart (not shown)to allow access for lay-up. These implementations are merely exemplary,and other implementations may also be used.

FIG. 21B illustrates frame plies on the mandrel in accordance with anembodiment of the present invention, as shown in FIG. 21A. As shown inFIG. 21B, frame plies 2010, may be formed and placed in a frame recess1620 on bag 1010, which sits atop media 1210, which sits atop armature1020.

FIG. 21C illustrates frame plies and a frame mandrel on the mandrel inaccordance with an embodiment of the present invention, as shown in FIG.21A. As shown in FIG. 21C, a frame mandrel, such as frame mandrel 1920is placed on top of frame plies 2010 on bag 1010, which sits atop media1210, which sits atop armature 1020. This implementation is merelyexemplary, and other implementations may also be used.

FIGS. 21A-21C have provided an overview of frames and frame mandrels.FIGS. 22-24B describe frame ply lay-up and frame mandrels in moredetail.

FIG. 22 illustrates wing attachment plies being applied to a mandrel toform wing attachment pockets in accordance with an embodiment of thepresent invention, as shown in FIGS. 21A-21C. As shown in FIG. 22, wingattachment plies 2050 have been formed and placed in wing attachmentpockets 2210. Wing attachment pockets 2210 may also include metalinserts (not shown). Metal inserts provide bearing strength in the jointareas.

FIG. 23 illustrates frame plies in frame recesses in a mandrel in moredetail in accordance with an embodiment of the present invention, asshown in FIGS. 21A-21C. As shown in FIG. 23, frame plies 2010 are placedinside frame recess 2310 in mandrel 1410. In this example, frame plies2010 are placed inside frame recess 2310 in the forward and lowerportion of access door opening 115. Frame mandrels 1920 may then beplaced on frame plies 2010, as shown in FIG. 24A.

FIGS. 24A-24B, illustrate the combination of frame plies and a framemandrel on a mandrel.

FIG. 24A illustrates a frame mandrel in a frame recess in a mandrel inmore detail in accordance with an embodiment of the present invention,as shown in FIGS. 21A-21C. As shown in FIG. 24A, frame mandrel 1920 isplaced in frame recess 1620 in mandrel 1610.

FIG. 24B illustrates a frame mandrel over frame plies in a frame recessin a mandrel in accordance with an embodiment of the present invention,as shown in FIGS. 21A-21C, 23, and 24A. As shown in FIG. 24B, frameplies 2010 are placed in frame recess 1620 (not shown). In this example,frame plies 2010 are placed inside frame recess 1620 in the forwardsection of access door opening 115. Frame mandrel 1920 is then placed ontop of frame plies 2010. As also shown in FIG. 24B, the recess in accessdoor 115 may be filled with a winding aid, such as, door opening fillerblock 2440. As described above, winding aids, such as door openingfiller block 2440, provide a convex surface for filament winding(described in the following paragraphs). As shown in FIG. 24B, a secondwinging aid 2450 is used to define ply cutting locations during cut anddrape 911 (described below).

FIGS. 21A-24A have described place frames and frame mandrels 908, asshown in FIG. 9. As shown in FIG. 9, following place frames and framemandrels 910, wind inner skin 910 occurs. Wind inner skin 910 isdescribed in FIGS. 25-28.

FIG. 25 illustrates preparing the mandrel for filament winding of theinner skin in accordance with an embodiment of the present invention, asshown in FIG. 9. As shown in FIG. 25, mandrel 1390 includes door recessfiller 1760, windshield recess filler 1750, and wing attachment pocketfillers 1740. Other frames and frame mandrels have also been inserted inmandrel 1390 (as described above.)

As further shown in FIG. 25, a hoop wrap 2510 may be applied as hoops2520 to mandrel 1390 by a filament winding machine (not shown here, butshown in FIG. 26). Hoop wrap 2510 is wound circumferentially by thefilament winding machine around mandrel 1390, such that space existsbetween the adjacent hoops 2520. In one implementation, approximately 4inches of “advance” is used. Advance is the space between subsequentwinding paths. In the case of a hoop wrap, advance is the distancebetween adjacent bands of fiber being placed by the winding machinehead. Hoop wrap 2510 is subsequently removed from mandrel 1390 aftersufficient filament of desired orientation has been applied to mandrel1390 to retain frame plies, frame mandrels, and winding aids.

FIG. 26 illustrates applying filament to the mandrel for filamentwinding of the inner skin by a filament winding machine in accordancewith an embodiment of the present invention, as shown in FIG. 25. Asshown in FIG. 26, a filament winding machine 2610 applies filament 2620(such as carbon fiber) to mandrel 1390. In the example shown in FIG. 26,approximately 15% of one internal ply is in place to form the innerskin. In one implementation, when the inner skin is complete across-section of the inner skin will be about 0.016 inch thick, overabout a 0.250 inch thick core. In this implementation, moreover, frameswill generally be about 0.034 inch thick with a height of about 1.25inch and a width of about 1.75 inch. These dimensions are provided forexemplary purposes and are typical of one fuselage structure for onetype of aircraft. Therefore, other implementations may be used, asneeded.

As shown in FIG. 26, filament winding machine 2610 is used to apply aninner skin to mandrel 1390. For filament winding of a structure (such asmandrel 1390), a filament winding machine having a capacity ofapproximately 25 feet in length and a swing of approximately 3 feet isadequate. An acceptable filament winding machine for this purpose iscommercially available through vendors, such as Entec in Salt Lake City,Utah. However, other filament winding machines may be used. In oneimplementation, the wind angle may be close to plus or minus 45 degrees,as practicable. For other implementations, it may be preferable to builda custom winding machine suited for a particular structure beingmanufactured. Also, in some implementations, during filament winding, itmay be necessary to stop and place doubler plies by hand, as needed.

FIG. 27A is a perspective view of a mandrel with a filament-wound innerskin in accordance with an embodiment of the present invention, as shownin FIGS. 21A-21C. As shown in FIG. 27A, mandrel 1610 has been fullywound with inner skin 2710 by filament winding machine 2610 (not shown).

FIG. 27B is a cut-away view of a mandrel with a filament-wound innerskin in accordance with the embodiment of the present invention, asshown in FIG. 27A. As shown in FIG. 27B, filament-wound inner skin 2710sits atop frame mandrel 1920, which sits atop frame plies 2010, whichsits atop bag 1010, which surrounds media 1210, which surrounds armature1020. Filament wound inner skin 2710 also covers winding aids 2720.

FIG. 28 is a side view of a mandrel with a filament-wound inner skinwith external end hoop plies in accordance with an embodiment of thepresent invention, as shown in FIG. 26. As shown in FIG. 28, externalend hoop plies 2810 are placed around portions of mandrel 1390 overinner skin 2820. External hoop plies 2810 are used to hold inner skin2810 for cut and drape 911 (described below).

FIGS. 25-28 have described wind inner skin 910, as shown in FIG. 9. Asshown in FIG. 9, following wind inner skin 910, cut and drape 911occurs. Cut and drape 911 is described in FIGS. 29-30B.

FIG. 29 illustrates cutting a mandrel in accordance with an embodimentof the present invention, as shown in FIG. 9. As shown in FIG. 29, ends2910 have been cut from inner skin 2820 over mandrel 1390. After ends2910 have been cut, the other portions of mandrel 1390 are cut (asdescribed below).

FIG. 30A is a perspective view of a mandrel with inner skin cut anddraped in accordance with an embodiment of the present invention, asshown in FIG. 27A. As shown in FIG. 30A, mandrel 1610 shows inner skin2710. During cut and drape 911 (as described in FIG. 9), inner skin 2710is cut in particular locations so that winding aids can be removed (asdescribed in FIG. 17). For example, as shown in FIG. 30A, winding aidsinclude door recess fillers and windshield area fillers. Other windingaids may also include passenger window recess fillers. Additionally,although not shown in FIG. 30A, winding aids 2450 may be used around thewinding aids to provide a cutting guide. These winding aids 2450identify the location of other winding aids. The winding aids 2450include raised pins, which may be used to position cutting aids forremoval of excess material from mandrel 1610.

FIG. 30B is a cut-away view of a mandrel with inner skin that has beencut and draped in accordance with an embodiment of the invention, asshown in FIG. 30A. As shown in FIG. 30B, after cutting inner skin 2710and removing winding aids in recess areas, joggle areas 3010 areexposed. Inner skin plies 2710 can now be draped into joggle areas 3010.Inner skin plies 2710 are draped into contact with frame plies 2010,frame mandrel 1920, and bag 1010, which surrounds media 1210, whichsurrounds armature 1020. Thus, following cutting and removal of windingaids (not shown) doubler plies 2020 (not shown) are draped to joggleareas 3010. Doubler plies 2020 are placed over inside corners where skinplies 2710 are cut to allow them to drape into joggled areas. Doublerplies 2020 reinforce the cut inner skin plies 2710. Joggled areas 3010are normally located around windshields, windows, and door openings.Joggled areas 3010 allow for the windows, the windshield, and the doorsto fit flush to the surface of the structure. Alternatively, joggledareas 3010 could be eliminated, where other solutions could be used tomake the flush fit. These implementations are merely exemplary, andother implementations may also be used.

FIGS. 29-30B have described cut and drape 911, as shown in FIG. 9. Asshown in FIG. 9, following cut and drape 911, place core 912 occurs.Machine core to shape 952 and place core 912 are described in FIGS.31A-32.

FIG. 31A illustrates machining core in accordance with an embodiment ofthe present invention, as shown in FIG. 9. As shown in FIG. 31A, coredetails are machined by cutting a desired peripheral shape from coresheet stock and then chamfering that periphery to provide sandwichmaterial for placement between inner and outer skins of the fuselage toenhance skin buckling strength. Core sheet stock may be prepared tothickness by a core material supplier or it may be cut to desiredthickness in a clean environment. Core materials include foam core aswell as honeycomb core materials. Foam core materials are made from hightemperature thermoplastics that have been foamed using a blowing agentor some other foaming methodology. Honeycomb core materials are madefrom metal foils or plastic materials (strengthened with natural orsynthetic fibers) formed into paper bonded together in such a manner asto resemble natural bee's wax honeycomb. Examples of plastic honeycombcore material include Nomex and Korex materials registered trademarks ofDupont. However, any type of core material may be used for machine coreto shape 952.

FIG. 31B is a perspective view of a mandrel with core material inaccordance with an embodiment of the present invention, as shown in FIG.9. As shown in FIG. 31B, mandrel 1610 includes core pieces 3110, whichare applied over film adhesive to the outside of inner skin 2710, whereinner skin 2710 has been draped into recesses in mandrel 1610. Corepieces 3110 are used to prevent skin buckling. Core pieces 3110 alsohelp to retain a desired structural shape.

FIG. 31C is a cut-away view of a mandrel with core details in accordancewith an embodiment of the present invention, as shown in FIG. 31A. Asshown in FIG. 31C, a core piece 3110 is placed in a recess created inmandrel 1610. Core piece 3110 sits atop inner skin 2710, which sits atopbag 1010, which sits atop media 1210, which sits atop armature 1020.Core pieces 3110 can be placed on inner skin plies 2710 with or withoutfilm adhesive. The use of film adhesive may be needed, if the windingresin being used does not have adhesive properties.

FIG. 32 illustrates a portion of a mandrel with film adhesive coveringcore material in accordance with an embodiment of the present invention,as shown in FIGS. 31A-31B. As shown in FIG. 32, core pieces 3110 includecore material with film adhesive 3210 and core material without adhesive(not shown). Separator film 3225 may also be placed on the inner skin.Separator film 3225, which is placed to aid manufacture, such as, whenit is time to remove excess material and to drape any joggle areas 3010.

FIGS. 31A-32 have described place core 912, as shown in FIG. 9. As shownin FIG. 9, following place core 912, wind outer skin 914 occurs. Windouter skin 914 is described in FIGS. 33-35.

FIG. 33 illustrates preparing a mandrel for application of an outer skinby a filament winding machine in accordance with an embodiment of thepresent invention, as shown in FIG. 9. As shown in FIG. 33, mandrel 1390includes end domes 1720, separator film 3225, and end hoop wraps 2810.End domes 1720 are used to provide a convex surface. Separator film 3225is used around the joggle areas to aid in removal of winding aids injoggle recesses. External end hoop wraps 2810 are used around mandrel1390 to hold inner skin plies 2630 in place. Further, as shown in FIG.33, filament for outer skin 3330 has begun to be applied to mandrel 1390by filament winding machine 2610. Further, hoop wrap 3340 may be appliedto mandrel 1390 by hand or by filament winding machine 2610. Hoop wrap3340 is wound circumferentially such that gaps exists between successivewraps but close enough together to hold assorted winding aids in theircorrect locations.

FIG. 34 illustrates applying an outer skin to a mandrel by a filamentwinding machine in accordance with an embodiment of the presentinvention, as shown in FIG. 33. As shown in FIG. 34, filament windingmachine 2610 wraps mandrel 1390 in filament for outer skin 3410. In thisexample, filament winding machine 2610 is at about in a 25% finishedstate, with outer skin 3410 applied to mandrel 1390. In oneimplementation, outer skin 3410 includes two plies. Alternatively, inanother implementation, filament winding machine 2610 may wind an outerskin 3410 that is twice as thick. Other implementations may have otherplies or other layers.

FIG. 35A is a perspective view of a mandrel with a filament wound outerskin in accordance with an embodiment of the present invention, as shownin FIG. 9. As shown in FIG. 35A, filament winding machine 2610 (notshown) applies outer skin 3510 to mandrel 1610. As also shown in FIG.35A, filament winding machine 2610 may apply an outer skin to a largearea, which may be larger than just a fuselage cabin (e.g., to include atail cone 106).

FIG. 35B is a cut-away view of a mandrel with a filament wound outerskin in accordance with an embodiment of the present invention, as shownin FIG. 35A. As shown in FIG. 35B, outer skin 3510 has been placed overcore pieces 3110, inner skin 2710, frame mandrel 1920, and other windingaids. Core piece 3110 sits atop inner skin 2710, which sits atop bag1010, which sits atop media 1210, which sits atop armature 1020.

As shown in FIG. 9, following wind outer skin 913, cut and drape 914occurs. Cut and drape 914 is described in FIGS. 36A-37.

FIG. 36A is a perspective view of a mandrel with outer skin cut anddraped in accordance with an embodiment of the present invention, asshown in FIG. 9. As shown in FIG. 36A, mandrel 1610 shows outer skin3510. During cut and drape 914 (as described in FIG. 9), outer skin 3510is cut in particular locations so that winding aids can be removed. Forexample, as shown in FIG. 36A, winding aids that are removed includewindshield winding aid 3610, door winding aids 3620, and window windingaid 3630.

FIG. 36B is a cut-away view of a mandrel with outer skin that has beencut and draped in accordance with an embodiment of the invention, asshown in FIG. 36A. As shown in FIG. 36B, after cutting outer skin 3510and removing winding aids in recess areas joggle area 3645 is exposed.Outer skin plies 3510 can now be draped into joggle area 3645. Outerskin plies are draped into contact with inner skin 2710, which hasalready been formed into joggle area 3645. Inner skin plies contactframe plies 2010, frame mandrel 1920 and bag 1010, which surrounds media1210, which surrounds armature 1020. Thus, following cutting and removalof winding aids (not shown) doubler plies 2020 (not shown) are draped tojoggle area 3645. Doubler plies 2020 are placed over inside cornerswhere outer skin plies 3510 are cut to allow them to drape into joggledareas. Doubler plies 2020 reinforce the cut outer skin plies 3510.Joggled areas 3645 are normally located around windshields, windows, anddoor openings. Joggled areas 3645 allow for the windows, the windshield,and the doors to fit flush to the surface of the structure.Alternatively, joggled areas 3645 could be eliminated, where othersolutions could be used to make the flush fit. For example, frames couldinclude joggle areas 3645, rather than using plies, such as doublerplies 2020.

FIG. 37 illustrates the mandrel after cutting and draping of the outerskin in accordance with an embodiment of the present invention, as shownin FIG. 33. As shown in FIG. 37, mandrel 1390 is shown with outer skin3410 cut and draped into door joggle area 3710. Separator film 3225 (notshown) has been removed from between inner skin 2630 and outer skin 3410to allow cutting and draping of outer skin 3410 into door joggle area3710. Separator film 3225 is any low cost thermoplastic film the mostprevalent being polyethylene and nylon. Other films may be usedincluding FEP, PTFE, and ECTFE. External end hoop plies 2810 areretaining outer skin 3410 and inner skin 2710 to mandrel 1390. End domes1720 can now be removed from both ends of mandrel 1390.

As shown in FIG. 9, following cut and drape 914, close mold 915 occurs.Close mold 915 is described in FIGS. 38A-39D.

FIG. 38A illustrates preparing a circumferential mold for a mandrel inaccordance with an embodiment of the present invention, as shown in FIG.9. As shown in FIG. 38, circumferential mold 3810 may include severalpieces (described below). Circumferential mold 3810 may be placed on theexterior of mandrel 1610. In one implementation, circumferential mold3810 is approximately 20 feet long, 4 feet wide, and 6 feet high. Inthis implementation, circumferential mold 3810 consists of three pieces:(1) lower circumferential mold section 3812, (2) left topcircumferential mold section 3814, and (3) right top circumferentialmold section 3816. In other implementations, circumferential mold 3810may be one piece, two pieces, or more than three pieces. Theseimplementations are merely exemplary, and other implementations may alsobe used.

FIG. 38B is a cut-away view of a mandrel in the circumferential mold inaccordance with an embodiment of the present invention. As shown in FIG.38B, circumferential mold 3810 is closed over mandrel 1610.Circumferential mold 3810 covers all portions of mandrel 1610, includingouter skin 3510 and outer joggle recesses 3645. Once circumferentialmold 3810 has been closed, a vacuum may be applied, so that all air isremoved between outer skin 3510 and inner skin 2710, between outer skin3510 and core piece 2060, between inner skin 2710 and bag 1010, andbetween inner skin 2710 and frame mandrel 1920, among other areas.Additionally, pressurization may be used with a vacuum. In thisimplementation, enclosed cavity 1030 containing media 1210 ispressurized. Between two and three atmospheres are generally adequatefor pressurization, although pressure may vary depending upon theparticular application. If pressurization is used, during cut and drape911 and/or cut and drape 914, cuts should be done to allow for expansionduring pressurization. In both of these implementations, the framemandrels 1920 may be placed under vacuum to maintain their shape(described below).

FIG. 39A illustrates preparing a circumferential mold with a vacuumsystem for the frame mandrels during curing in accordance with anembodiment of the present invention, as shown in FIGS. 38A-38B. As shownin FIG. 39A, vacuum system 3910 may be used so that frame mandrels (notshown here, but shown in FIG. 39B) maintain the proper shape. For vacuumsystem 3910, internal plumbing (not shown) is needed. Standard vacuumplumbing may accomplish these tasks.

FIG. 39B illustrates a cut-away of the mandrel in the circumferentialmold with a vacuum system for the frame mandrels in accordance with anembodiment of the present invention, as shown in FIG. 39A. Vacuum system3910 provides for pulling a vacuum, using piping which starts at endplates (not shown), continues through media 1210, and goes to bag 1010,which then goes to a vacuum port 3920. Thus, a vacuum is transmitted tothe frame mandrels, such as frame mandrel 1920, through vacuum port3920, which may be installed when the frame mandrels are positioned inmandrel 1610.

FIG. 39C illustrates a vacuum port in a frame mandrel in accordance withan embodiment of the present invention, as shown in FIGS. 39A and 39B.As shown in FIG. 39C, vacuum port 3920 is a couple between the interiorof frame mandrel 1920 and vacuum source tube 3970. Vacuum from vacuumsource tube 3970 is extended into the interior of frame mandrel 1920,which has media 1210 inside, through double-ended needle 3925.Double-ended needle 3925 passes through valve 3965 in frame mandrel 1920and valve 3965 in mandrel 1610. Because frame mandrel 1920 is filledwith media 1210 it is necessary to equip frame mandrel with filteringdevice 3960 to prevent media 1210 from plugging double ended needle3925.

FIG. 39D illustrates a device for maintaining a vacuum in a framemandrel in accordance with an embodiment of the present invention, asshown in FIGS. 39B and 39C. As shown in FIG. 39C, vacuum port 3920includes a double ended inflation needle 3925, of such a length that oneend of needle 3925 when inserted into a valve 3965 in frame mandrel 1920extends into air space 3962 and the other end inserted into a valve 3965in mandrel 1610 extends into vacuum source tube 3970. Further, needle3925 is modified at ends 3927, 3928 with side holes 3922 to prevent endplugging and further comprises a disk 3924 located approximately at themid-point of needle 3925 to guard against end plugging by sealing end3927 against the filtering device 3960.

As shown in FIG. 9, following close mold 915, cure 916 occurs. Cure 916is described in FIG. 40.

FIG. 40 illustrates curing a filament wound mandrel in a circumferentialmold in an oven in accordance with an embodiment of the presentinvention, as shown in FIG. 9. As shown in FIG. 40, in oneimplementation, circumferential mold 3810 is placed in oven 4010. Inthis implementation, oven heat cures the composite materials on mandrel1390 against circumferential mold 3810. Alternatively, heat can beapplied using integral heating methods, such as circulating heatedliquid through tubes within circumferential mold 3810. Alternatively,heat can also be used inside enclosed cavity 1030 to cure the compositematerial. Indeed, any type of oven or any type of heat can be used tocure composite material inside circumferential mold 3810.

As shown in FIG. 9, following cure 916, de-mold 917 occurs. De-mold 917is described in FIGS. 41-45. De-mold 917 includes removingcircumferential mold 3810, as described in FIG. 41, removing media 1210,as described in FIG. 42, removing armature 1020, as described in FIG.43, removing bag 1010, as described in FIG. 44, and making bag 1020available for reuse, as described in FIG. 45.

FIG. 41 illustrates removing a circumferential mold from around aone-piece integrally stiffened fuselage on a mandrel in accordance withan embodiment of the present invention, as shown in FIG. 9. As shown inFIG. 41, fuselage 4110 is removed from circumferential mold 3810. In oneimplementation, removal of circumferential mold 3810 occurs afterfuselage 4110 has been allowed to cool sufficiently, for example, tobelow 150° F. For this implementation after mandrel 160 has cooled, avacuum and/or pressure is also released. Other implementations may beused.

FIG. 42 illustrates removing media from a mandrel in accordance with anembodiment of the present invention, as shown in FIG. 9. As shown inFIG. 42, mandrel 1390 is inside one-piece integrally stiffened fuselagecabin, which in turn is inside circumferential mold 3810. In oneimplementation, a vacuum 4210 removes media through fill ports (notshown) in end plates 1380 (also not shown). In this implementation,after removal of media from mandrel 1390, armature 1320 is removed frommandrel 1390 (not shown), and armature 1320 may then be reused. Otherimplementations may be also used.

FIG. 43 illustrates a one-piece integrally stiffened fuselage containedin a circumferential mold after removal of media and armature inaccordance with one embodiment of the present invention as shown in FIG.42. As shown in FIG. 43, media and armature 1320 (not shown) have beenremoved from mandrel 1390.

FIG. 44 illustrates removing a bag from a one-piece integrally stiffenedfuselage in accordance with an embodiment of the present invention asshown in FIG. 41. As shown in FIG. 44, bag 1010 is removed fromone-piece integrally stiffened fuselage 4110, while both are supportedon work stands 4410. Bag 1010 as previously explained has no substantialshape without armature 1020 and media 1210.

FIG. 45 illustrates a bag after removal from a mandrel in accordancewith an embodiment of the present invention, as shown in FIG. 44. Bag1010 may now be reused, after removal from one-piece integrallystiffened fuselage 4110.

As shown in FIG. 9, following de-mold 917, extract frame mandrels 918occurs. Extract frame mandrels 918 is described in FIG. 46.

FIG. 46 illustrates removing frame mandrels from a one-piece integrallystiffened fuselage in accordance with an embodiment of the presentinvention as shown in FIG. 9. As shown in FIG. 46, once media has beenextracted, and once the bag 1420 has been extracted, frame mandrels 1920(not shown) may then be extracted. Once frame mandrels 1920 have beenextracted from mandrel 4200, a structure can be seen, such as, in thisexample, one-piece integrally stiffened fuselage 4610. Fuselage 4610includes door openings 4610, windshield opening 4620, and doorattachment points 4630. Fuselage 4610 also depicts other components, asshown in FIG. 46.

As shown in FIG. 9, following extract frame mandrels 918, visuallyinspect parts 919, trim 920, and prime and paint 921 occurs. Oneimplementation of actions 919, 920, and 921 is described in FIGS. 47-48.Other implementations may be used.

FIG. 47 illustrates a one-piece integrally-stiffened fuselagemanufactured in accordance with one embodiment of the present inventionas shown in FIG. 9. As shown in FIG. 47, fuselage 4700 has been preparedfor inspection. During visually inspecting part 919, fuselage 4700 isexamined visually, both interior and exterior surfaces. In addition,visually inspect part 919 includes verification that dimensionaltolerances are correct. In addition to visually inspect part 919, trim920 includes trimming any material, as needed. Trim 920 includes manualmethods (such as a hand held air powered router motor with router tool)or automatic methods (such as robot using a router tool). In addition totrim 920, prime and paint 921 includes sanding and filling surfaces toan acceptable level of smoothness. After sanding and filling, fuselage4700 receives paint primer on all exterior surfaces.

FIG. 48 illustrates a one-piece integrally stiffened fuselagemanufactured in accordance with one embodiment of the present inventionas shown in FIG. 9. As shown in FIG. 48, fuselage 4800 includes both afuselage cabin 4810 and a tail cone 4820. FIG. 48 demonstrates thatfuselage 4800 includes both fuselage cabin 4810 and tail cone 4820. Inother implementations, other parts of an aircraft may further beincluded with fuselage 4800.

As shown in FIG. 9, following prime and paint 921, store for assembly922 occurs. Store for assembly involves storing a structure, such as afuselage, until needed, e.g., until the fuselage is needed to assemblean aircraft.

As described in the preceding sections, various implementations may beused to create a structure, such as fuselage 4700 or fuselage 4800. Thefollowing section illustrates one of many such alternativeimplementations.

2. Alternate 2

FIG. 49 is a block diagram illustrating the process of manufacturing aone-piece fuselage in accordance with another embodiment of the presentinvention, as shown in FIG. 9. As shown in FIG. 49, system 4900 issubstantially similar to system 900 shown in FIG. 9. However, as shownin FIG. 49, close composite mold 4906 has replaced close mold 915,prepare autoclave 4904 has replaced prepare oven 975, and cure usingautoclave 4902 has replaced cure 916. Close composite mold 4906 involvesthe use of a mold manufactured from composite materials. In oneimplementation, the mold is manufactured from high temperaturefiber-reinforced plastics. However, in other implementations, othermolds and other materials could be used. Prepare autoclave 4904 and cureusing autoclave show that an autoclave is used to cure the structure insystem 4900. Cure using autoclave pressure is described below. Closecomposite mold 4906, prepare autoclave 4904, and cure using autoclave4902 are described in FIGS. 50-53.

FIG. 50 illustrates assembling a circumferential mold around a mandrelin accordance with an embodiment of the present invention, as shown inFIG. 49. As shown in FIG. 50, fuselage 5000 may be prepared for curingaccording to prepare autoclave 4904 (from FIG. 49). Fuselage 5000 may beplaced in a mold 5010 for curing. In one implementation, mold 5010consists of seven pieces. Four of the seven pieces are visible in FIG.50: (1) upper left half 5002, (2) upper right half 5004, (3) lowerforward segment 5008, and (4) lower aft segment 5006. The other segments(not shown) comprise: (1) left windshield area 5012 (not shown), (2) theright windshield area 5014 (not shown), and (3) the bulkhead flangeupper half segment 5016 (not shown). In other implementations, more orless pieces could be used to construct the mold These implementationsare merely exemplary, and other implementations may also be used.

FIG. 51 illustrates bagging a circumferential mold in accordance with anembodiment of the present invention, as shown in FIG. 50. As shown inFIG. 51, the seven pieces of mold 5010 (as described in FIG. 50) havebeen placed around fuselage 5000. After construction of the mold 5010,it is placed in a bag 5102 for sealing to form bag assembly 5100. Bag5102 may be made of nylon. However, bag 5102 may also be made of anyother material. In one implementation, bag 5102 is sealed such that avacuum is formed between the bag 1420 and envelope bag 5102. In thisimplementation, the vacuum provides for a void-free structure. Otherimplementations may also be used.

FIG. 52 illustrates placing a circumferential mold in an autoclave forcuring in accordance with an embodiment of the present invention, asshown in FIG. 51. As shown in FIG. 52, bag assembly 5200 may be placedin an autoclave 5202. In one implementation, autoclave 5202 appliespressure to bag assembly 5200. In this implementation, during autoclavecuring, the frames and frame mandrels are maintained under by vacuum tomaintain the proper shape for the frames. In this implementation,further curing by autoclave 5202 generally takes 1½ to 2 hours at atemperature between 250° F. and 350° F. at one to three atmospheres ofpressure. The pressure is applied to the outside of bag 5102 and to themedia cavity between armature 1320 and bag 1420. Autoclaves such asthose manufactured by Thermal Equipment Corporation, TariccoCorporation, McGill AirPressure Corporation, Melco Steel Incorporated,and American Autoclave Company may be used. These implementations areonly exemplary, and other implementations and other types of autoclavesmay also be used.

FIG. 53 illustrates removing a circumferential mold after curing in anautoclave in accordance with an embodiment of the present invention, asshown in FIG. 52. As shown in FIG. 53, the pieces of mold 5010 areremoved following curing by autoclave 5202. In FIG. 53, upper left half5002 is removed, while upper right half 5004 is still in place.

As shown in FIG. 49, following de-mold 917, extract mandrels 918,visually inspect part 919, trim 920, and prime and paint 921 occurs. Oneimplementation of actions 919, 920, and 921 is depicted in FIG. 54.Other implementations may be used.

FIG. 54 illustrates a one-piece integrally-stiffened fuselagemanufactured in accordance with another embodiment of the presentinvention, as shown in FIG. 49. As shown in FIG. 54, fuselage 5400 hasbeen prepared for inspection. During visually inspect parts 919,fuselage 5400 is examined visually, both interior and exterior surfaces.In addition, visually inspect parts 919 includes verification thatdimensional tolerances are correct. In addition to visually inspectparts 919, trim 920 includes trimming any material, as needed. Trim 920includes manual methods (such as a hand held air powered router motorwith router tool) or automatic methods (such as robot using a routertool). In addition to trim 920, prime and paint 921 includes sanding andfilling surfaces to an acceptable level of smoothness. After sanding andfilling, fuselage 5400 receives paint primer on all exterior surfaces.

3. Other Alternates:

Alternates 1 (such as process alternate #1805) and Alternate 2 (such asprocess alternate #2815) are described herein, but any number ofalternate methods and structures are possible for a one-piece structure,such as a fuselage, using the claimed invention.

VI. CONCLUSION

As described above, therefore, other embodiments of the invention willbe apparent to those skilled in the art from consideration of thespecification and practice of the invention disclosed herein. It isintended that the specification and examples be considered as exemplaryonly, with a true scope and spirit of the invention being indicated bythe following claims and their equivalents. In this context, equivalentsmean each and every implementation for carrying out the functionsrecited in the claims, even if not explicitly described therein.

What is claimed is:
 1. A method of manufacturing a one-piece closed-shape structure using a mandrel, comprising: preparing the mandrel, wherein the mandrel comprises a bag and an armature; applying a frame mandrel to the mandrel to form a frame for the structure; filling the mandrel and the frame mandrel with media; applying a curable resign to a fiber; applying the fiber over the mandrel and frame mandrel to form the structure; curing the structure; removing the media from the mandrel and frame mandrel; and extracting the mandrel and frame mandrel from the structure.
 2. The method of claim 1, wherein preparing further comprises: placing the armature through the bag; and conforming the shape of the bag to a desired shape of the structure.
 3. The method of claim 2, wherein conforming further comprises: sealing the bag; placing the armature and the bag in a form tool; and conforming the shape of the bag to the form tool.
 4. The method of claim 3, wherein conforming further comprises: filling a space between the armature and the bag with air; and creating a vacuum between the form tool and the bag to force the bag to conform to the shape of the form tool.
 5. The method of claim 1, wherein applying a frame mandrel further comprises: applying a frame ply to an exterior of the bag; and applying the frame mandrel over the frame ply.
 6. The method of claim 1, wherein filling further comprises compacting the media.
 7. The method of claim 6, wherein compacting further comprises vibrating the mandrel and frame mandrel to aid compaction.
 8. The method of claim 1, wherein applying the fiber comprises: winding the fiber over the mandrel and frame mandrel to form the structure.
 9. The method of claim 8, wherein winding further comprises: placing a first winding aid on the bag; winding the fiber over the first winding aid, the frame mandrel, and the mandrel to form an inner skin; cutting the inner skin to remove the first winding aids; placing a second winding aid on the inner skin; winding the fiber over the second winding aid and inner skin to form an outer skin; and cutting the outer skin to remove the second winding aids.
 10. The method of claim 9, wherein placing second winding aids further comprises placing a core piece on the inner skin.
 11. The method of claim 1, wherein curing further comprises: placing a mold around an exterior of the structure; sealing the mold; placing the mold in a heating device; and applying heat to the mold using the heating device.
 12. The method of claim 11, wherein curing further comprises: creating a vacuum in the mandrel; and creating a vacuum in the frame mandrel.
 13. The method of claim 1, wherein curing further comprises: placing a mold around an exterior of the structure; sealing the mold; placing the mold in an autoclave; and applying pressure to the mold using the autoclave.
 14. The method of claim 1, wherein the structure is a fuselage of an aircraft.
 15. A computer-implemented method of manufacturing a one-piece closed-shape structure using a mandrel, comprising: preparing the mandrel, wherein the mandrel comprises a bag and an armature; applying a frame mandrel to the mandrel to form a frame for the structure; filling the mandrel and the frame mandrel with media; applying a curable resign to a fiber; applying the fiber over the mandrel and frame mandrel to form the structure; curing the structure; removing the media from the mandrel and frame mandrel; and extracting the mandrel and frame mandrel from the structure.
 16. A method of manufacturing a one-piece closed-shape structure, using a mandrel comprising: preparing the mandrel, wherein the mandrel comprises a bag and an armature; placing the armature through the bag; conforming the shape of the bag to a desired shape of the structure; applying a frame mandrel to the mandrel to form a frame of the structure; filling the mandrel and the frame mandrel with media; applying a curable resign to a fiber; applying the fiber over the frame mandrel and the bag to form an inner skin; placing a core piece on the inner skin; applying the fiber over the core piece and inner skin to form an outer skin; placing a mold around an exterior of the structure; curing the structure in the mold; removing the mold from the structure; removing the media from the mandrel and the mandrel frame; extracting the armature from the bag; and extracting the bag from the structure.
 17. The method of claim 16, wherein conforming further comprises: sealing the bag; placing the armature and the bag in a form tool; and conforming the shape of the bag to the form tool.
 18. The method of claim 17, wherein conforming further comprises: filling a space between the armature and the bag with air; and creating a vacuum between the form tool and the bag to force the bag to conform to the shape of the form tool.
 19. The method of claim 16, wherein applying a frame mandrel further comprises: applying a frame ply to an exterior of the bag; and applying a frame mandrel over the frame ply.
 20. The method of claim 16, wherein filling further comprises compacting the media.
 21. The method of claim 20, wherein compacting further comprises vibrating the mandrel and frame mandrel to aid compaction.
 22. The method of claim 16, wherein applying the fiber over the frame mandrel and the bag to form an inner skin comprises: winding the fiber over the frame mandrel and the bag to form the inner skin.
 23. The method of claim 22, wherein winding further comprises: placing a winding aid on the bag; winding the fiber over the frame mandrels, the winding aid, and the bag to form the inner skin; and cutting the inner skin to remove the winding aid.
 24. The method of claim 16, wherein applying the fiber over the core piece and inner skin to form an outer skin comprises: winding the fiber over the core piece and inner skin to form the outer skin.
 25. The method of claim 24, wherein winding further comprises: placing a winding aid on the inner skin; winding the fiber over the core piece, the winding aid, and the inner skin to form an outer skin; and cutting the outer skin to remove the winding aid.
 26. The method of claim 16, wherein curing further comprises: sealing the mold; placing the mold in a heating device; and applying heat to the mold using the heating device.
 27. The method of claim 26, wherein curing further comprises: creating a vacuum in the mandrel; and creating a vacuum in the frame mandrel.
 28. The method of claim 16, wherein curing further comprises: sealing the mold; placing the mold in an autoclave; and applying pressure to the mold using the autoclave.
 29. The method of claim 16, wherein the one-piece closed-shape structure is an airplane fuselage.
 30. A computer-implemented method of manufacturing a one-piece closed-shape structure, using a mandrel comprising: preparing the mandrel, wherein the mandrel comprises a bag and an armature; placing the armature through the bag; conforming the shape of the bag to a desired shape of the structure; applying a frame mandrel to the mandrel to form a frame of the structure; filling the mandrel and the frame mandrel with media; applying a curable resign to a fiber; applying the fiber over the frame mandrel and the bag to form an inner skin; placing a core piece on the inner skin; applying the fiber over the core piece and inner skin to form an outer skin; placing a mold around an exterior of the structure; curing the structure in the mold; removing the mold from the structure; removing the media from the mandrel and the mandrel frame; extracting the armature from the bag; and extracting the bag from the structure. 